1. Field of the Invention
The present invention relates generally to an industrial gas turbine engine, and more specifically to a low cooling flow rotor blade for an industrial gas turbine engine.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine, such as an industrial gas turbine (IGT) engine, includes a turbine section with multiple rows or stages of rotor blades that react with a high temperature gas flow to convert the energy of the gas flow into mechanical energy that is used to drive the engine and power an aircraft in the case of an aero engine or drive an electric generator in the case of an IGT engine. The engine efficiency can be increased by passing a higher temperature gas flow into the turbine. However, the turbine inlet temperature is limited by the material properties of the turbine parts, especially the first stage stator vanes and rotor blades, and to the amount of cooling that these parts can be provided with.
In future engines, higher turbine inlet temperatures will require improved materials and better cooling designs. Improvements in the cooling capabilities of the airfoils (vanes and blades) will also contribute to improved engine efficiencies. Since the pressurized cooling air used to cool the airfoils is typically bled off from the compressor, the work performed to pressurize the cooling air is lost in the engine since the cooling air does no work on the engine. Low cooling flow vanes and blades are thus desired in order to provide improved efficiencies.
FIG. 1 shows an external heat transfer coefficient profile for a prior art first stage turbine rotor blade. As indicated by the figure, a suction side immediately downstream of the leading edge, as well as a pressure side trailing edge region of the airfoil, is exposed to higher hot gas side external heat transfer coefficient than the mid-chord section of the pressure side and downstream of the suction side surfaces. The graph shows (point A) a high Q on the suction side, a high heat load on the aft section of the pressure side surface (point C), and a high heat load region for the blade leading edge (point B) on this graph. In general, the heat load for the airfoil aft section is higher than in the forward section.
As the TBC technology improves and more IGT turbine blades are applied with a thicker or low conductivity TBC, the cooling air flow demand is gradually being reduced. As a result, there is not sufficient cooling flow for the design to split the total cooling flow into three flow circuits that utilize the forward flowing serpentine cooling design. Cooling flow for the blade leading edge and trailing edge has to be combined with a mid-chord flow circuit to form a single 5-pass flow circuit. However, for the forward 5-pass flow circuit with total blade cooling flow BFM (back flow margin) may become a design issue. In addition, a single 5-pass aft flowing serpentine for a large chord blade design may yield too high of a cooling air temperature when the cooling air reaches the end of the serpentine cooling flow channel. This results in a loss of cooling potential for the cooling air to achieve a designed low metal temperature. A low metal temperature for an airfoil is desired in order to increase the part life and to minimize erosion of the part.